1. Field of the Invention
This invention relates to rocket exhaust control apparatus and, more particularly, to apparatus designed to diminish the undesirable heat transfer and pressure effects resulting from rocket exhaust impingement.
2. Description of the Related Art
The thermal and pressure impingement effects of exhaust gases emanating from a rocket motor nozzle extend longitudinally well beyond the exit plane of the nozzle because of the concentrated supersonic flow pattern characteristic of the exhaust gas plume. Relatively small tactical missile rocket motors have nozzle exit velocities on the order of Mach 3.5 with recovery pressures of 300 pounds per square inch and recovery temperatures in excess of 5,000 degrees Fahrenheit. These high velocities, pressures and temperatures encountered in tactical or test firings of the rocket motor remain concentrated in the plume, extending its length, and generally result in destructive final pressure and thermal impingement effects in a relatively small area to the rear of the motor. In open environment static test firings, this necessitates rugged, thermally protected structures positioned a considerable distance to the rear of the rocket nozzle, or a relatively vast open area behind the rocket. Test firings of the rocket motor into a duct or plenum require that similar features be incorporated in the plenum design to avoid destruction of the enclosure by the exhaust gas impingement. In tactical firings, the launching pad and components must also be protected. Provision of structures that can withstand the adverse temperature and pressure of the rocket motor exhaust increases the complexity and cost of the installations involved if firing safety is to be preserved.
Apparatus for reducing the pressure and thermal impingement effects of a supersonic exhaust gas plume emanating from a rocket motor nozzle is disclosed in my prior U.S. Pat. No. 4,733,751, entitled "ROCKET EXHAUST DISRUPTER", the contents of which are incorporated herein by reference. That patent discloses an exhaust tube and duct system having a disrupter body positioned within the exhaust tube along the central longitudinal axis thereof to disrupt and disperse the plume concentration. The body is generally cylindrical in shape and has a central passage through which a portion of the exhaust may pass. The body and the passage are sized and shaped to provide a flow area which is normal to the exhaust flow and less than the cross-sectional area of the exhaust plume at the position of the body. The reduced flow area causes the exhaust flow to undergo a normal Shock wave upstream of the body before passing through and around the body at a subsonic velocity. This resulting reduced flow velocity significantly reduces the downstream pressure and thermal effects of the exhaust gas plume.
A similar disrupter member is installed in a dissipator device which is the subject of my prior U.S. Pat. No. 4,709,780, entitled "EXHAUST DISSIPATOR/DISRUPTER DEVICE". The dissipator/disrupter device disclosed in that patent comprises a container having an inlet for connection to a rocket exhaust with a plurality of exhaust orifices to disperse the exhaust flow out of the container in different directions. A disrupter device like that described hereinabove is mounted within the container more or less like the disrupter device in the exhaust tube of U.S. Pat. No. 4,733,751.
The principle of disrupting the rocket exhaust flow by disrupter devices such as those disclosed in the two patents referenced hereinabove is based upon the discovery that the undisturbed supersonic rocket exhaust impingement heat transfer and recovery pressure effects can be substantially reduced if the Mach number of the exhaust, at the location of impingement, can be significantly lowered. The disrupter devices of my two prior patents are disclosed as generally cylindrical shapes with tapered bores which totally enclose at least a portion of the rocket exhaust at a particular axial location in the exhaust plenum. The present invention is concerned with particularly shaped devices which are arrayed in positions relative to the rocket exhaust which are significantly different from the tapered cylinder disrupters previously disclosed. These shapes are not hollow, are not cylindrical and do not enclose any portion of the exhaust plume.
I am aware of the following patents which generally relate to the subject of controlling or affecting jet engine exhaust gas streams U.S. Pat. Nos.:
3,174,282 Harrison PA1 3,611,726 Medawar PA1 3,655,007 Hillbig PA1 3,706,353 Ffowces-Williams et al PA1 3,708,036 Duthion et al PA1 3,820,626 Bonneaud et al PA1 3,976,160 Hoch et al PA1 2,702,986 Kadosch et al PA1 3,434,666 Shaw PA1 3,739,872 McNair PA1 4,007,587 Banthin et al
The first seven of these patents deal with noise suppressors for jet engines. The next three pertain to thrust reversers, gas deflectors or defusers. The last patent in this group relates to suppression of infrared radiation from a jet engine. All of these patents relate to jet engine exhausts; none is concerned with rocket exhausts which are much more intense and present different problems from jet engine exhausts. Finally, none of these patents indicates any appreciation of developing a normal shock wave in an exhaust plume to reduce the gas flow to subsonic velocity.